Particular combustion systems for gas turbine engines utilize combustors which burn a gaseous or liquid fuel mixed with compressed air. Generally, a combustor includes a fuel nozzle assembly including multiple fuel nozzles which extend downstream from an end cover of the combustor and which provide a mixture of fuel and compressed air to a primary combustion zone or chamber. A liner or sleeve circumferentially surrounds a portion of the fuel nozzle assembly and may at least partially define the primary combustion chamber. The liner may at least partially define a hot gas path for routing combustion gases from the primary combustion zone to an inlet of a turbine of the gas turbine.
In operation, compressed air flows through a premix or swozzle portion of each fuel nozzle. Fuel is injected into the compressed air flow and premixes with the compressed air before it is routed into the combustion chamber and burned to produce the combustion gases. During operation, various operating parameters such as fuel temperature, fuel composition, ambient operating conditions and/or operational load on the gas turbine may result in combustion dynamics or pressure pulses within the combustor. The combustion dynamics may cause oscillation of the various combustor hardware components such as the liner and/or the premix fuel nozzle which may result in undesirable wear of those components.